Efficiency of a gas turbine increases as the temperature of the combustor outlet or the temperature of the turbine inlet increases. However, the temperature of the combustor outlet of gas turbines in current use reaches 1,500° C., and the temperature of the gas turbine blade surface which is exposed to high-temperature combustion gas exceeds the critical temperature of the heat-resistant alloy used. Therefore, it is necessary to cool the gas turbine blades.
To do so, compressed air is provided from a compressor to a cooling pass formed inside the gas turbine blade and convection cooling of the cooling pass wall takes place. Also, film cooling is performed in such a way that a plurality of through-holes are provided on the surface of the gas turbine blade and air is ejected therethrough from the cooling pass onto the surface of the gas turbine blade and flown on the entire surface. Thus, the increase in the temperature of the gas turbine blade is suppressed to a point lower than the critical temperature.
With regard to the film cooling structure, elliptically-shaped holes have been proposed (e.g., patent literature 1 and patent literature 2) with the intention of forming a layer of cooling air on the entire surface of the gas turbine blade.